1. Field of the Invention
The present invention relates generally to gas turbine engine nozzles and, more particularly, to a divergent section of a nozzle and its cooling means to reduce the engine's infrared signature.
2. Discussion of the Background Art
Hot nozzles emit infrared radiation (IR) which is highly undesirable for military combat aircraft. Infrared radiation from gas turbine engines is conventionally suppressed by shielding and cooling the hot metal structures of the engine. The use of outer flaps and seals around the divergent section of variable nozzles is demonstrated in U.S. Pat. No. 4,128,208, by Ryan et al., entitled "Exhaust Nozzle Flap Seal Arrangement", assigned to the same assignee as the present invention. Nozzles also require cooling for structural reasons. Cooling air is conventionally drawn from the fan section or a compressor section of the gas turbine engine which is expensive in terms of fuel and power consumption. Ejecting nozzles such as the type used on some General Electric J79 engine models have employed slot type ejectors to induct ambient cooling air from the atmosphere to supplement the engine supplied cooling air in order to reduce the use of the more expensive engine air. Such ejecting nozzles provided efficient cooling for variable nozzle throats but are not as effective for cooling thrust vectoring nozzles such as axisymmetric vectoring exhaust nozzles which have 2 DOF pivoting flaps and seals. U.S. patent application Ser. No. 07/700,979, entitled "AXISYMMETRIC VECTORING EXHAUST NOZZLE THERMAL SHIELD", disclosed a shielding means and a nozzle cooling means for axisymmetric vectoring nozzles to efficiently cool the nozzle and shield it from emitting infrared radiation. However, faster cooling means are required when an emergency situation occurs while the nozzle is operating hot.
One type of conventional gas turbine engine exhaust nozzle includes primary and secondary exhaust flaps arranged for defining a variable area convergent-divergent exhaust nozzle. The exhaust nozzle is generally axisymmetric, or annular, and exhaust flow is confined by the primary or convergent flaps and secondary or divergent flaps being positioned circumferentially adjacent to each other, respectively.
The divergent flaps, for example, have a forward end defining a throat of minimum flow area and an aft end having a larger flow area for defining a diverging nozzle extending downstream from the throat. The divergent flaps are variable, which means that the spacing between the divergent flaps as they are moved from a smaller radius position to a larger radius position must necessarily increase. Accordingly, conventional exhaust nozzle seals are suitably secured between adjacent divergent flaps to confine the exhaust flow and prevent leakage of exhaust flow between the divergent flaps.
An advanced axisymmetric vectoring nozzle has been developed and patented in U.S. Pat. No. 4,994,660, entitled "AXISYMMETRIC VECTORING EXHAUST NOZZLE", by Hauer, assigned to the present assignee, and herein incorporated by reference. An axisymmetric vectoring nozzle provides a means for vectoring the thrust of an axisymmetric convergent/divergent nozzle by universally pivoting the divergent or divergent flaps of the nozzle in an asymmetric fashion or in other words pivoting the divergent flaps in radial and tangential directions with respect to the unvectored nozzle centerline.
Vectoring nozzles, and in particular axisymmetric vectoring exhaust nozzles of the type disclosed in the Hauer reference, provide positionable divergent flaps. These divergent flaps are positionable not only symmetrically, relative to a longitudinal centerline of the exhaust nozzle, but may also are positionable asymmetrically relative thereto for obtaining pitch and yaw vectoring. An exemplary thrust vectoring nozzle uses three vectoring actuators to translate and tilt a vectoring ring which in turn forces the divergent flaps in predetermined positions. The vectoring ring tilt angle and tilt direction establish the nozzle's vector angle and vector direction, respectively. Axial translation of the vectoring ring establishes the exit area (often referred to as A9) for a given throat area (often referred to as A8).
Retaining the flaps and adjacent seals in place for an axisymmetric nozzle is very difficult because of the varying degree of askewness between the flaps and seals encountered during asymmetric operation of the nozzle for thrust vectoring. The seal has to be retained radially, with respect to the nozzle's centerline, and circumferentially to prevent the flap seals from becoming unfeathered from the adjacent flaps. Radial retention means between seals and flaps is disclosed in U.S. Pat. No. 5,269,467, entitled "Vectoring Exhaust Nozzle Seal and Flap Retaining Apparatus", and was developed to counteract the inverse exhaust pressures that occur when there is higher pressure on the radially outer surfaces of the seal and flap than on the radially inner surfaces.
Modern multi-mission aircraft application employ engines, such as the GE F110 engine, with convergent/divergent nozzles to meet operational requirements. Convergent/divergent nozzles have, in serial flow relationship, a convergent section, a throat, and a divergent section. Characteristically, these nozzles employ variable area means at both the nozzle throat and at the nozzle exit. This provides a means to maintain a desired exit to throat area ratio which in turn allows efficient control over the operation of the nozzle. The operation of the nozzle is designed to provide a nozzle exit/throat area (A9/A8) schedule which is optimized for the design cycle of the engine and ideally should provide efficient control at both low subsonic and high supersonic flight conditions. These types of nozzles typically use pneumatic or hydraulic actuators to provide the variable operation. Typically, the exit and throat areas are mechanically coupled to each other in such a manner as to create an area ratio (A9/A8) schedule which is a function of nozzle throat area (A8). The area ratio schedule is typically predetermined to provide efficient engine operation across a wide range of engine conditions but typically optimum performance at specific engine conditions is compromised somewhat in order to provide adequate efficiency throughout the range of engine operation. Thrust vectoring nozzles typically have the ability to independently control nozzle exit area and throat area which allows the engine to achieve a higher level of performance across a wide range of engine operating conditions. An additional benefit of independent throat and exit area control is the capability to overexpand the nozzle divergent system beyond its optimal performance area ratio to create divergent system static wall pressure lower than ambient pressures to thereby pull the lower temperature ambient air into the nozzle where it can be used to cool the divergent system components. Ambient pressures, as with all ambient conditions, refer to freestream conditions outside the aircraft. Ambient conditions are also generally found in unpressurized nozzle bays, i.e. the area surrounding the convergent and divergent flaps that lie inside outer flaps or engine or other casings surrounding the nozzle flaps.
The successful operation of combat aircraft is dependent, in part, upon the ability of the aircraft to remain undetected by infrared sensors of various ground and air based weapon systems, such as ground and air launched missiles, during flight. The high temperatures of the engine's exhaust gases and the hot metal turbine parts and the hot metal walls directly in contact with the hot gases cause the engine to emit high levels of infrared energy. Military aircraft engaged in combat are vulnerable to anti-aircraft missiles employing highly sophisticated infrared sensors.
A number of apparatus have been designed to reduce infrared emissions from gas turbine engines. Each type of design endeavors to provide a combination of aerodynamics, heat transfer, and geometry which will result in an effective IR suppressor for the least suppressor weight and power effects on a turbine engine. One of these types of geometries utilizes a concentric centerbody within an annular duct. This suppressor geometry is referred to as a plug or centerbody suppressor and exemplified by U.S. Pat. Nos. 4,214,441, 4,044,555, 3,970,252 and the like. The plug suppressors are supported and fed cooling air from fan and/or high pressure bleed air aerodynamically shaped struts which are also used to position and support the centerbody. These hollow centerbody plug suppressors consume expensive fan and compressor air engine and power and result in reduced engine efficiency and combat operating radius.
There exists a need for a means to provide rapid emergency mode cooling of the interior hot surfaces of the divergent section of the nozzle to rapidly suppress the engine's IR signature during combat missions with a minimal adverse effect on the overall operability of the aircraft and its engine.